Dual fan gas turbine engine and gear train

ABSTRACT

A gas turbine engine includes a first fan, a second fan spaced axially from the first fan, a turbine-driven fan shaft and an epicyclic gear train coupled to be driven by the turbine-driven fan shaft and coupled to drive the first fan and the second fan. The epicyclic gear train includes a carrier that supports star gears that mesh with a sun gear, and a ring gear that surrounds and meshes with the star gears. The star gears are supported on respective journal bearings.

CROSS REFERENCE TO RELATED APPLICATIONS

The present disclosure is a continuation in part of U.S. patentapplication Ser. No. 11/906,602, filed Oct. 3, 2007.

BACKGROUND OF THE INVENTION

The present invention is generally directed to gas turbine engines and,more particularly, to gear systems for use in variable cycle gas turbineengines. Typical turbofan gas turbine engines utilize a fan to generatetwo streams of air for producing thrust. The fan pushes a first streamof inlet air into a core turbine engine where the inlet air is used tosustain a combustion process. The first stream of inlet air is passedthrough a series of compressors, a combustor and turbines, which aredisposed in an axially serial relationship. The compressors increase thedensity and temperature of the air for carrying out the combustionprocess in the combustor. High-energy gases resulting from thecombustion process are then used to produce thrust and rotate theturbines. Typically, a first turbine is used to drive the compressorsfor the combustion process, and a second turbine is used to rotate thefan. The fan also pushes a second stream of inlet air around the coreengine to directly produce thrust. Typically, each engine is configuredwith a fixed bypass duct that dictates the engine bypass ratio: thedistribution of the inlet air routed to the core engine (primary air)and the inlet air routed around the core engine (bypass air).

Typically, turbofan engines are configured to operate optimally at oneset of operating conditions, the design point of the engine. Forexample, many small military aircraft are configured with low bypassratio engines, where rapid engine response time and high thrust aredesirable. Conversely, high bypass ratio engines achieve lower noiseemissions and better fuel efficiency and are therefore well suited forlarge transportation aircraft. Thus, the design point of a high bypassratio turbofan is typically configured to operate at cruising conditionssuch that low thrust specific fuel consumption is achieved. At operatingconditions above this design point, such as high-thrust takeoffconditions, specific fuel consumption increases as the engine requiresmore fuel for marginal thrust increases. Below this design point,specific fuel consumption also decreases as a disproportionate amount ofthrust is lost as fuel usage is scaled back. These types of turbofansare sometimes referred to as single cycle engines, as the operatingperformance of the engine is tied to one specific thermodynamic cyclerelating the air flow to the combustion process. Specifically, a singlecycle engine is limited by the mass flow rate of the inlet air producedby the fan at any given rotational speed. Thus, an engine designer mustchoose between efficiency and performance in selecting the design pointfor a single cycle engine.

The operating range of a turbofan engine, and hence flexibility in theperformance of the aircraft in which it is used, can be increased byvarying the bypass ratio. Variable cycle engines operate in multiplemodes, each with a different thermodynamic cycle in which the mass flowrates of each mode are selected to meet different performance needs. Forexample, two-cycle engines operate in either a high bypass configurationor a low bypass configuration, in which a variable bypass duct istypically used to divert inlet air from the bypass duct to the coreengine. With the added flexibility, however, comes added complexity inmatching air flow speeds, pressure ratios, mass flow rates, blade speedsand the like. For example, in order to accommodate the variable bypassduct, two-cycle engines frequently utilize two-stage fans having a largediameter forward section and a small diameter aft section. The forwardsection is thus able to push inlet air into both the bypass duct and thecore engine. In one configuration, however, these fan sections areconnected to the same input turbine shaft. Thus, the speeds of each fansection cannot be individually controlled and blade tip speed controldifficulties arise. In another configuration, the forward fan isconnected to the low pressure turbine shaft and the rear fan isconnected to the high pressure turbine shaft. Thus, the speed of eachfan is governed by overall engine performance concerns rather thanbypass ratio and mass flow rate concerns. There is, therefore, a needfor a variable cycle fan section that permits individualized controlover the rotational speeds of different fan sections.

SUMMARY OF THE INVENTION

A disclosed example turbine engine according to an exemplary embodimentincludes a first fan, a second fan spaced axially from the first fan, aturbine-driven fan shaft, and an epicyclic gear train coupled to bedriven by the turbine-driven fan shaft and coupled to drive the firstfan and the second fan. The epicyclic gear train including a carriersupporting star gears that mesh with a sun gear, and a ring gearsurrounding and meshing with the star gears, the star gears beingsupported on respective journal bearings.

In a further embodiment of the foregoing turbine engine, the epicyclicgear train is coupled to drive the first fan and the second fan atdifferent rotational speeds.

In a further embodiment of the foregoing turbine engine, the epicyclicgear train is coupled to directly drive the second fan without arotational speed change relative to the turbine-driven fan shaft andcoupled to drive the first fan at a reduced rotational speed relative tothe turbine-driven fan shaft.

In a further embodiment of the foregoing turbine engine, the first fanis coupled to be driven from the carrier and the second fan is coupledto be driven from the ring gear.

In a further embodiment of the foregoing turbine engine, the epicyclicgear train has a gear reduction ratio such thatωSecond-Stage=ωFirst-Stage·[(NR+Ns)/Ns], where ωFirst-Stage is arotational speed of the first fan, ωSecond-Stage is a rotational speedof the second fan, NR is the number of teeth on the ring gear and NS isthe number of teeth on the sun gear.

In a further embodiment of the foregoing turbine engine, the epicyclicgear train has a gear reduction ratio of greater than or equal to about2.3.

In a further embodiment of the foregoing turbine engine, the epicyclicgear train has a gear reduction ratio of greater than or equal to 2.3.

In a further embodiment of the foregoing turbine engine, the epicyclicgear train has a gear reduction ratio of greater than or equal to about2.5.

In a further embodiment of the foregoing turbine engine the epicyclicgear train has a gear reduction ratio of greater than or equal to 2.5.

In a further embodiment of the foregoing turbine engine, the epicyclicgear train has a gear reduction ratio of about 1.37:1.

In a further embodiment of the foregoing turbine engine, the first fanand the second fan define a bypass ratio of greater than about ten (10)with regard to a bypass airflow and a core airflow.

In a further embodiment of the foregoing turbine engine, the first fanand the second fan define a bypass ratio of greater than about 10.5:1with regard to a bypass airflow and a core airflow.

In a further embodiment of the foregoing turbine engine, the first fanand the second fan define a bypass ratio of greater than ten (10) withregard to a bypass airflow and a core airflow.

In a further embodiment of the foregoing turbine engine, wherein thefirst fan and the second fan define an overall fan pressure ratio thatis less than about 1.45.

In a further embodiment of the foregoing turbine engine, the first fanand the second fan define an overall fan pressure ratio that is that isless than 1.45.

Although different examples have the specific components shown in theillustrations, embodiments of this invention are not limited to thoseparticular combinations. It is possible to use some of the components orfeatures from one of the examples in combination with features orcomponents of another of the examples.

These and other features disclosed herein can be best understood fromthe following specification and drawings, the following of which is abrief description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a partial cross-sectional view of a front portion of a gasturbine engine illustrating a turbo fan, epicyclic gear train and acompressor section.

FIG. 2 is an enlarged cross-sectional view of the epicyclic gear trainshown in FIG. 1.

FIG. 3 is an enlarged cross-sectional view of an example ring gearsimilar to the arrangement shown in FIG. 2.

FIG. 4 is a view of the ring gear shown in FIG. 3 viewed in a directionthat faces the teeth of the ring gear in FIG. 3.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT

A portion of a gas turbine engine 10 is shown schematically in FIG. 1.The turbine engine 10 includes a fixed housing 12 that is constructedfrom numerous pieces secured to one another. A compressor section 14having compressor hubs 16 with blades are driven by a turbine shaft 25about an axis A. A turbo fan 18 is supported on a turbo fan shaft 20that is driven by a compressor shaft 24, which supports the compressorhubs 16, through an epicyclic gear train 22.

FIG. 1 shows a cross sectional view of the propulsive section ofvariable cycle turbofan engine 10, which includes two-stage fan section12 configured for delivering multiple streams of inlet air I to turbofanengine 10 in order to vary the bypass ratio of engine 10. Fan section 12is configured to generate bypass air for producing thrust, and togenerate core air for compressing in compressor 13. The core air is usedto sustain a combustion process within a core portion (not shown) ofengine 10 positioned axially downstream of compressor 13. Two-stage fansection 12 includes first-stage blade 14A and second-stage blade 14Bthat are inter-disposed between inlet guide vane (IGV) 16, intermediateguide vane 18 and exit guide vane 20. First-stage blade 14A andsecond-stage blade 14B are joined at their inner diameter ends to fanshaft system 22, and vanes 16, 18 and 20 are joined at their outerdiameter ends to engine housing 24. Fan shaft system 22 comprises fanshaft 26, first-stage fan shaft 28, second-stage fan shaft 30 andepicyclic gear system 32. Engine housing 24 comprises inlet case 34,gate 36, fan case 38 and intermediate duct 40. Fan section 12 and fanshaft system 22 comprise radially symmetric systems that are configuredfor rotating about axis A in FIG. 1. Thus, fan section 12 includescircular arrays of first-stage blades 14A, second-stage blades 14B andvanes 16, 18 and 20, which are mounted to annular shafting members 26,28 and 30.

Inlet case 34 and fan case 38 are coaxially aligned to form outer bypassduct 44 and main duct 46 for initially splitting inlet air I into firststream I1 and second stream I2. First and second streams I1 and I2 aredirected into intermediate duct 40, whereby second stream I2 is furtherparsed to produce third stream I3. Gate 36 is selectively opened andclosed, for example, by an engine control system such as a FullAuthority Digital Engine Control (FADEC) (not illustrated), to vary theamount of inlet air I that is permitted to bypass compressor 13. Thus,the amount of thrust directly produced by fan section 12 in proportionto the amount of thrust produced by the core engine changes.Accordingly, the performance of engine 10 can be optimized over a widerange of operating conditions. To further maximize the efficiency andperformance of engine 10, fan shaft system 22 is provided with epicyclicgear system 32 to coordinate the rotational speeds of first-stage blades14A and second-stage blades 14B to produce the required mass flow rateof inlet air I as is needed by engine 10 while avoiding blade tip speedissues.

In the high bypass ratio configuration of engine 10, inlet air I entersinlet case 34 and is directed into outer bypass duct 44 and main duct 46by IGV 16. Inner annular duct 48, which is provided in segments alonginlet case 34, gate 36, first-stage blade 14A and fan case 38, dividesinlet air I into first and second streams I1 and I2. First stream I1 isdriven by tip-blade 50 of first-stage blade 14A to intermediate duct 40.Second stream I2 is driven by first-stage blade 14A and second-stageblade 14B within main duct 46 to intermediate duct 40. After passingthrough fan section 12, first stream I, and second stream I2 continuepast the core section of engine 10 and are expelled from engine 10through an exhaust nozzle (not shown) to produce thrust. In variousembodiments, second stream I2 is reunited with first stream I1 forpassing out of engine 10 at the exhaust nozzle. Intermediate duct 40,however, also includes flow deflector 52 that extracts third stream I3from second stream I2 for supplying to compressor 13. Compressor 13compresses third stream I3 such that it can be used in a combustionprocess to generate high-energy gases for turning turbines that drivecompressor 13 and fan section 12.

In order to operate engine 10 in the low bypass ratio configuration,such as to increase the proportion of inlet air I provided to compressor13, gate 36 is actuated by the FADEC to close-off bypass duct 44. Gate36 comprises a pivotable door that can be rotated to deflect firststream I, from bypass duct 44 to main duct 46 to increase the mass flowrates of second stream I, and third stream I3. Due to the differentoperating conditions of engine 10, different demands are placed onfirst-stage blade 14A and second-stage blade 14B. First-stage blade 14Ais configured for driving first stream I1, second stream I2 and thirdstream I3, while second-stage blade 14B is only configured for drivingsecond stream I2 and third stream I3. First-stage blade 14A is radiallylonger and generally sturdier than second-stage blade 14B such thatfirst stream I1 can be driven via blades 50. Thus, first-stage blade 14Agenerates higher tip speed and increased mass flow rates thansecond-stage blade 14B if rotated at the same speed. It is, however,desirable for both blades to operate near, but below, their respectivecritical tip speeds. It is also desirable to match the mass flow ratesgenerated by first-stage blade 14A and second-stage blade 14B such thatsurge, stall or other undesirable engine conditions are avoided withinfan section 12. Fan shaft system 22 is therefore provided with epicyclicgear system 32 to individually optimize the rotational speeds offirst-stage blade 14A and second-stage blade 14B. Specifically,epicyclic gear system 32 reduces the rotational speed of first-stageblade 14A such that tip-blades 50 do not exceed their critical speed,and first-stage blades 14A do not provide more inlet air I than can bemoved by second-stage blades 14B.

Fan shaft 26 is connected with compressor 13 through compressor shaft54, which is connected with a turbine in the core of engine 10. In oneembodiment, compressor 13 is connected to a low pressure turbinedisposed axially rearward of a high pressure turbine within the coreengine. Compressor shaft 54 is supported within engine 10 by strut 56,which is stationary and connected with intermediate case 40 and bearing58. In one embodiment of engine 10, strut 56 is connected with an oildispensing system (not shown) to conduct oil to bearing 58 forlubrication and cooling purposes. Fan shaft 26 extends from compressorshaft 54 and connects with second-stage shaft 30 at connection 60.Second-stage shaft 30 and fan shaft 26 include radially extendingflanges that permit additional components of engine 10 and gear system32 to be secured between second-stage shaft 30 and fan shaft 26.Second-stage shaft 30 is joined with second-stage rotor 62, which drivessecond-stage blade 14B. Thus, second-stage blade 14B is directly andrigidly connected with fan shaft 26. Fan shaft 26 and second-stage shaft30 are supported within engine 10 by strut 64, which extends from exitguide vane 20 to bearings 66A and 66B. Connection 60 also links fanshaft 26 to epicyclic gear system 32, which is also connected tofirst-stage shaft 28. First-stage shaft 28 connects with first-stagerotor 64, which drives first-stage blade 14A. Thus, first-stage blade14A is connected to compressor shaft 54 through epicyclic gear system 32and fan shaft 26. First-stage shaft 28 is supported within second-stagerotor 62 by bearing 68 and bearing 69. Stationary torque tube 70 extendsfrom epicyclic gear system 32 to inlet guide vane 16, whereby stationarytorque tube 70 is anchored to fan inlet case 34 to provide a staticmember for supporting epicyclic gear system 32 and to counteract torqueloads. Epicyclic gear system 32 comprises a speed-reducing gear systemthat reduces the rotational speed of first-stage blade 14A to a speedbelow that of fan shaft 26 and second-stage blade 14B. Accordingly,first-stage blades 14A and second-stage blades 14B rotate at speedsclose to their respective tip speed limits. Also, first-stage blades 14Aand second-stage blades 14B rotate at speeds that coordinate efficientmass flow of inlet air I through bypass duct 44 and main duct 46 of fansection 12 over a wide range of engine performance parameters.

FIG. 2 shows a schematic cross section of epicyclic gear system 32disposed within engine 10 of FIG. 1. Engine 10 includes fan shaft 26,first-stage shaft 28, second-stage shaft 30, epicyclic gear system 32,compressor shaft 54, connection 60, second-stage rotor 62, strut 64,bearings 66A and 66B, bearing 68, and static torque tube 70, as wereseen in FIG. 1. Engine 10 also includes oil feed 72, oil manifold 74,oil sump 76 and oil drain 78. Epicyclic gear system 32 comprises ringgear 80, planet gears 82, sun gear 84, journal bearing 86 and carrier88.

Compressor shaft 54 extends axially forward from the turbine and rotatesabout centerline A. Fan shaft 26 is connected to the forward end ofcompressor shaft 54 through, for example, a spline connection at firstend 90. Second end 92 of fan shaft 26 is connected to first end 94 ofsecond-stage shaft 28 at connection 60. Second end 96 is connected withsecond-stage rotor 62 with, for example, a spline connection at secondend 96. Together, fan shaft 26 and second-stage shaft 30 form a bulgedor arched shaft that connects compressor shaft 54 with second-stagerotor 62. Generally, the bulged or arched portion of shaft formed by fanshaft 26 and second-stage shaft 30 is positioned at flanged connection60, where epicyclic gear system 32 is connected to fan shaft 26 andsecond-stage shaft 30. Fan shaft 26 extends axially forward and radiallyoutward from compressor shaft 54, while first end 94 of second-stageshaft 30 extends axially forward and radially inward from second end 92of fan shaft 26. Second end 96 of second-stage shaft 30 comprises alarger diameter than first end 90 of fan shaft 26, and second end 92 offan shaft 26 comprises a larger diameter than second end 96 ofsecond-stage shaft 30. As such, the fan shaft 26 and second-stage shaft30 provide a means for fan shaft 26 to extend past epicyclic gear system32 in a radially compact manner, while still permitting first-stageshaft 28 and static torque tube 70 to connect with epicyclic gear system32.

Static torque tube 70 extends coaxially about axis A from a stationarymember within engine 10, such as illustrated in FIG. 1, such that torquetube 70 provides a non-rotating member upon which to anchor gear system32. Thus, gear system 32 is suspended within engine 10 between statictorque tube 70 and shafts 26 and 30. First-stage shaft 28 extends fromgear system 32 between torque tube 70 and second-stage shaft 30, pastsecond-stage rotor 62, to support first-stage rotor 64. Compressor shaft54 is rotated by a turbine located within the core of gas turbine engine10. As the core gas turbine engine rotates compressor shaft 54, fanshaft 26 and second-stage shaft 30 rotate second-stage rotor 62 at thesame speed. Through the gear reduction of epicyclic gear system 32,first-stage shaft 28 rotates first-stage rotor 64 about axis A at aslower rate than the rate at which second-stage shaft 30 rotates aboutaxis A.

First end 90 of fan shaft 26 is supported within engine 10 by bearing66A, which rides on wing 98 of strut 64. Second end 96 of second-stageshaft 30 is supported by bearing 66B, which rides on wing 100 of strut64. Bearing 68 supports first-stage shaft 28 within second-stage shaft30. Bearing 68 is configured to react the radial and thrust loads fromfirst stage fan blades 14A, and transmits these loads to second-stagefan shaft 30 and fan shaft 26. In the embodiment shown, bearing 66Acomprises a tapered roller bearing, bearing 66B comprises a rollerbearing and bearing 68 comprises a tapered roller bearing, but in otherembodiments of the invention; tapered, roller or ball bearings may beused as needed in various designs.

FIG. 3 shows a schematic, front view of epicyclic gear system 32 of FIG.2. FIG. 3 shows the plurality of planet gears 82 situated within carrier88 as would be taken at section 3-3 of FIG. 2. FIG. 4 shows an explodedview of epicyclic gear system 32 of FIG. 2. When viewed in conjunctionwith each other, FIGS. 1 through 4 illustrate the lubrication systemincluded with engine 10 to provide oil to epicyclic gear system 32.Static torque tube 70 is stationary within engine 10 such that oil feed72 is readily extended through static torque tube 70 to provide oil toepicyclic gear system 32. The forward end of oil feed 72 is connected toan oil dispensing system (not shown), while the downstream end of oilfeed 72 is connected with universal joint 102 that connects with oilmanifold 74. Oil manifold 74 is connected with carrier 88 with, forexample, fasteners 104. Universal joint 102 allows oil manifold 74 torotate with carrier 88, while oil feed 72 remains static. Lubricatingoil travels through oil feed 72 and universal joint 102 and, throughcentrifical force, into oil manifold 74. Oil manifold 74 includesfittings 106 that dispense the lubricating oil into internal cavities108 within journal bearings 86. Journal bearings 86 include radiallyextending passages 110 that permit the lubricating oil to enter theinterfaces between planet gears 82 and journal bearings 86. Thelubrication oil thereafter axially spreads across the journal/gearinterface and is directed by a deflector (not shown) within carrier 88to the interfaces of planet gears 82 with ring gear 80 and sun gear 84by carrier 88. Additionally, as seen in FIG. 3, epicyclic gear system 32is provided with oil baffles 111 that force lubrication oil into theteeth of gears 80, 82 and 84. Baffles 111 are configured to ride incircumferential grooves 112 (FIG. 4) within planet gears 82. Thus,during operation of engine 10, oil is centrifically pushed throughepicyclic gear system 32 to provide lubrication and cooling of gears 80,82 and 84. U.S. Pat. No. 6,223,616 by Sheridan, which is hereinincorporated by this reference, explains in greater detail a lubricationsystem suitable for use in another embodiment of epicyclic gear system32 of the present invention.

Engine 10 is also provided with an oil collection system and emergencyoil distribution system. Wings 98 and 100 are connected to strut 64 withthreaded fastener 113, and come together to form oil sump 76 aboveconnection 60. Wing 98 includes an opening such that oil drain 78 hasaccess to the area between wing 98 and fan shaft 26. Knife edge seals114A and 114B are bolted to flanged connection 60 such that seal 114B isabove second-stage shaft 30 and seal 114A is above fan shaft 26.Deflectors 116A and 116B extend from wings 98 and 100, respectively, toabut seals 114A and 116B, respectively. As such, a passageway is formedbetween connection 60 and oil drain 78. Lubricating oil is pushedthrough epicyclic gear system 32 and toward connection 60 throughcentrifical operation of engine 10. The lubrication oil is permitted toseep through connection 60 and into sump 76 whereby oil drain 78collects the used lubrication oil. In one embodiment, oil drain 78 isconnected with a two-way pump, which collects the dispensed oil. Actingin suction mode, the pump collects the dispersed lubrication fromepicyclic gear system 32, whereby the oil is routed through filters andreturned to various oil dispensing systems within engine 10, such asthose used in conjunction with strut 56 (FIG. 1), or oil feed 72. In anemergency situation, the pump can be operated in pump mode to dispensethe lubrication oil from a storage tank, such as used in conjunctionwith strut 56 or oil feed 72, to sump 76 and forced into epicyclic gearsystem 32. Thus, epicyclic gear system 32 includes means for permittinglubrication oil to be circulated through gears 80, 82 and 84 utilizingvarious systems located within engine 10.

When viewed collectively, FIGS. 1 through 4 also illustrate the gearreduction achieved by epicyclic gear system 32. As noted above, statictorque tube 70 is fixedly disposed within engine 10. Sun gear 84 isrigidly affixed to first end 118 of torque tube 70 such that sun gear 84is also non-rotating and coaxial with axis A. Sun gear 84 is, in oneembodiment, force fit onto torque tube 70. Gear carrier 88 includes acentral bore into which sun gear 84 and static torque tube 70 areconcentrically disposed. Planet gear 82 is disposed within gear slot 120of gear carrier 88 and comprises one of a plurality of planet gearscircumferentially arranged within gear slots of gear carrier 88. Gearcarrier 88 comprises a pair of plates positioned on the forward and aftsides of planet gears 82 to provide a carriage for synchronizing therotation and maintaining spacing of the plurality of planet gears 82about sun gear 84 and axis A. A plurality of journal pins 86 extendthrough gear slots 120 and provide an axis upon which the plurality ofplanet gears 82 rotate. Journal pins 86 include internal cavities 108and radial passages 110 that receive a lubrication fluid from oilmanifold 74. Ring gear 80 is disposed concentrically about gear carrier88 to engage planet gears 82. In the embodiment shown, ring gear 80comprises forward and aft portions that are bolted together atconnection 60 through a flanged coupling to form an annular body thatrotates about axis A. Planet gears 82 include gear teeth that mesh withgear teeth of planet gear 82 and sun gear 84. In one embodiment, ringgear 80, sun gear 84 and planet gears 82 include double-helical, orherringbone, tooth configurations that facilitate flow of lubricationoil through epicyclic gear system 32.

Ring gear 80 of epicyclic gear system 32 is connected to fan shaft 26and second stage shaft 30 through flanged connection 60. Fan shaft 26 isconnected to compressor shaft 54 such that fan shaft drives ring gear 80at the same speed as compressor shaft 54. Likewise, second-stage shaft30 is connected to second-stage rotor 62 to drive second-stage fanblades 14B at the same speed as compressor shaft 54. Thus, second-stagefan blades 14B are directly driven by compressor shaft 54 withoutneeding to go through a gear reduction system. Ring gear 80 is alsodirectly driven by compressor shaft 54 to drive first-stage shaft 28through planet gears 82 and sun gear 84. First-stage shaft 28 isconnected to carrier 88 with, for example, threaded fasteners at bores122 of carrier 88 and bores 124 of first-stage shaft 28. In otherembodiments, second-stage shaft 28 is connected to carrier 88 using aflexible torque plate coupling in conjunction with bores 122, as isdescribed in U.S. Pat. No. 5,466,198 by McKibbin et al., which is hereinincorporated by this reference. As compressor shaft rotates ring gear80, ring gear 80 spins planet gears 82 around stationary sun gear 84.Planet gears 82 pull gear carrier 88 about axis A in the same rotationaldirection as ring gear 80, but at a reduced rate. Thus, first stageshaft 28 is driven to rotate about axis A in the same direction assecond-stage shaft 30, but at a slower speed.

Epicyclic gear system 32 achieves a reduction in the rotational speed offirst-stage shaft 28 commensurate with optimal operation of two-stagefan section 12. For example, differential gear systems, in which the sungear, the carrier and the ring gear all rotate, have been employed torotate counter-rotating fan sections in other gas turbine propulsionsystems. Differential gear reduction systems, however, achieve gearreduction ratios on the order of 8:1 or more, which would result in thefirst-stage fan shaft rotating at about thirteen percent of that of thesecond-stage fan shaft. In co-rotating, two-stage turbofans, it isdesirable that the two fan stages rotate at similar speeds such thatmass flows generated by each stage are closely matched. With the gearsystem of the present invention, in which sun gear 84 is maintainedstationary by static torque tube 70, lower gear reduction ratios areattainable, as is illustrated by equation (1), where NR is the number ofteeth in ring gear 80 and Ns is the number of teeth in sun gear 84.

(NR+Ns)/NR  Equation (1)

In one embodiment of epicyclic gear system 32, the ring gear has 117teeth, the planet gears have 37 teeth, and the sun gear has 43 teethsuch that a gear reduction ratio of about 1.37:1 is achieved. Thus, therotational speed of first-stage fan shaft 28 is reduced to aboutseventy-three percent of that of second-stage fan shaft 30, as isillustrated by equation (2), where ω_(Second-Stage) is the rotationalspeed of second-stage shaft 30 and ω_(First-Stage) is the rotationalspeed of first-stage shaft 28.

ω_(Second-Stage)=ω_(First-Stage)·[(NR+Ns)/Ns]  Equation (2)

In another embodiment of the invention, the rotational speed offirst-stage fan shaft 28 is reduced to about eighty-three percent ofthat of second-stage fan shaft 30.

In one disclosed, non-limiting embodiment, the engine 10 has a bypassratio that is greater than about six (6) to ten (10), the epicyclic geartrain 22 is a planetary gear system or other gear system with a gearreduction ratio of greater than about 2.3 or greater than about 2.5, anda low pressure turbine of the engine 10 has a pressure ratio that isgreater than about 5. In one disclosed embodiment, the engine 10 bypassratio is greater than about ten (10:1) or greater than 10.5:1, thediameter of at least one of the fans 14A or 14B is significantly largerthan that of the low pressure compressor 13, and the low pressureturbine has a pressure ratio that is greater than 5:1. In one example,the epicyclic gear train 22 has a gear reduction ratio of greater thanabout 2.3:1 or greater than about 2.5:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by a bypass flow B due to thehigh bypass ratio. The fans of the engine 10 are designed for aparticular flight condition—typically cruise at about 0.8M and about35,000 feet. The flight condition of 0.8 M and 35,000 ft, with theengine at its best fuel consumption—also known as “bucket cruiseTSFC”—is the industry standard parameter of 1 bm of fuel being burneddivided by 1 bf of thrust the engine produces at that minimum point.“Low fan pressure ratio” is the pressure ratio across the fan bladealone. The low overall fan pressure ratio across the first and secondfans as disclosed herein according to one non-limiting embodiment isless than about 1.45. “Low corrected fan tip speed” is the actual fantip speed in ft/sec divided by an industry standard temperaturecorrection of [(Tambient deg R)/518.7)̂0.5]. The “Low corrected fan tipspeed” as disclosed herein according to one non-limiting embodiment isless than about 1150 ft/second.

Additionally, epicyclic gear system 32 achieves more efficient torquetransfer from compressor shaft 54 to first-stage shaft 28 andsecond-stage shaft 30. Typical counter-rotating fan gas turbinepropulsion systems utilize a differential gear system that transferspower from a single rotational input to two rotational outputs. Thus,both power outputs are derived from the gear system, which producesinefficiencies in the power transfer and requires more robust gearsystems. Epicyclic gear system 32 of the present invention, however,permits power from compressor shaft 54 to be directly transferred tosecond-stage shaft 30. Second-stage shaft 30 is directly coupled tocompressor shaft 54 through fan shaft 26. Thus, gear reductioninefficiencies only result in transferring power to first-stage shaft 28from epicyclic gear system 32. Also, since only first-stage shaft 28receives power from gear system 32, gear system 32 need only besufficiently robust to transfer a portion of the power of compressorshaft 54, rather than all of the power of compressor shaft 54. This alsopermits epicyclic gear system 32 to achieve smaller diameters and morecompact designs.

Although the present invention has been described with reference topreferred embodiments, workers skilled in the art will recognize thatchanges may be made in form and detail without departing from the spiritand scope of the invention.

1. A turbine engine comprising: a first fan; a second fan spaced axiallyfrom the first fan; a turbine-driven fan shaft; and an epicyclic geartrain coupled to be driven by the turbine-driven fan shaft and coupledto drive the first fan and the second fan, the epicyclic gear trainincluding a carrier supporting star gears that mesh with a sun gear, anda ring gear surrounding and meshing with the star gears, the star gearsbeing supported on respective journal bearings.
 2. The turbine engine asrecited in claim 1, wherein the epicyclic gear train is coupled to drivethe first fan and the second fan at different rotational speeds.
 3. Theturbine engine as recited in claim 1, wherein the epicyclic gear trainis coupled to directly drive the second fan without a rotational speedchange relative to the turbine-driven fan shaft and coupled to drive thefirst fan at a reduced rotational speed relative to the turbine-drivenfan shaft.
 4. The turbine engine as recited in claim 1, wherein thefirst fan is coupled to be driven from the carrier and the second fan iscoupled to be driven from the ring gear.
 5. The turbine engine asrecited in claim 1, wherein the epicyclic gear train has a gearreduction ratio such that ω_(Second-Stage)=ω_(First-Stage)·[(NR+Ns)/Ns],where ω_(First-Stage) is a rotational speed of the first fan,ω_(Second-Stage) is a rotational speed of the second fan, NR is thenumber of teeth on the ring gear and NS is the number of teeth on thesun gear.
 6. The turbine engine as recited in claim 1, wherein theepicyclic gear train has a gear reduction ratio of greater than or equalto about 2.3.
 7. The turbine engine as recited in claim 1, wherein theepicyclic gear train has a gear reduction ratio of greater than or equalto 2.3.
 8. The turbine engine as recited in claim 1, wherein theepicyclic gear train has a gear reduction ratio of greater than or equalto about 2.5.
 9. The turbine engine as recited in claim 1, wherein theepicyclic gear train has a gear reduction ratio of greater than or equalto 2.5.
 10. The turbine engine as recited in claim 1, wherein theepicyclic gear train has a gear reduction ratio of about 1.37:1.
 11. Theturbine engine as recited in claim 1, wherein the first fan and thesecond fan define a bypass ratio of greater than about ten (10) withregard to a bypass airflow and a core airflow.
 12. The turbine engine asrecited in claim 1, wherein the first fan and the second fan define abypass ratio of greater than about 10.5:1 with regard to a bypassairflow and a core airflow.
 13. The turbine engine as recited in claim1, wherein the first fan and the second fan define a bypass ratio ofgreater than ten (10) with regard to a bypass airflow and a coreairflow.
 14. The turbine engine as recited in claim 1, wherein the firstfan and the second fan define an overall fan pressure ratio that is lessthan about 1.45.
 15. The turbine engine as recited in claim 1, whereinthe first fan and the second fan define an overall fan pressure ratiothat is that is less than 1.45.